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The strength of the aircraft

The principle of safe damages. Safety of aircraft is directly related to durability.

The design is called a safe operation, if required minimum inspection and repairs at a satisfactory core functions. Satisfactory performance is negligible probability of structural failure for a civil aircraft or acceptably low probability of failure for military aircraft. Safety of passengers and crew of civil aircraft is of paramount importance. Methods of analysis of structures, reliable in operation, designed primarily for civilian aircraft.

Modern plane is semi-monocoque type structure consisting of thin-walled sheets, supported by beams (farms) and the stringers to prevent buckling. The outer skin or wall forms an aerodynamic contour of the unit - the fuselage, wing, stabilizer. The stiffeners are attached to the inner surface of the skin and perceive concentrated loads. This design for many years served as the main object of aerodynamic research and distinguishes devices from conventional building designs.

The required service life of a civil aviation aircraft is determined on the basis of comprehensive economic considerations. It is 10-15 years old. The designer first of all tries to ensure a longer operation of the aircraft without the formation of cracks. To do this, he uses the developed calculation technique, which minimizes the stress concentration and tries to keep the voltage at the lowest possible level, based on the requirements for flight performance. For parts that are difficult to repair or replace, the designer may attempt to providedesired durability without cracking, equal to the lifetime of the aircraft. For many designs it is impossible. In addition, there is a risk of damage to structures serving transport, stumbling on the runway and decaying parts of the engine or the propeller. The designer must minimize the loss of strength resulting in occurrence of cracks or fatigue damage during operation of the aircraft. He solves this problem as follows:

picks up the material and determine the dimensions of parts to ensure adequate structural strength in cracks;

uses elements of safety (track variable loads and traffic, hindering the development of cracks);

selects a material having a low rate of fatigue cracks.

One of the modern means of improving the reliability of designs while increasing resource reduction of materials and improve economic efficiency - the design and definition of the duration of operation on the principle of safe damages. This takes into account the presence of structural elements of the initial metallurgical and technological defects and the formation of cracks in them the accumulation of operational damage.

The development and implementation of the principle of safe damage is possible only with the application of methods of fracture mechanics. The determination of the stress-strain state of structural elements containing defects such as cracks is the most responsible and complex step in calculating strength. In accordance with generally accepted ideas, the stress-strain state of a body with a crack is fully characterized by the values ​​of the stress intensity factor. On their preliminary definition, practically all the currently known criteria for brittle and quasi-brittle fracture are based, as well as the dependencies describing the growth of fatigue cracks.

The concept of "safe damage" refers to a design designed in such a way as to minimize the possibility of an airplane failure due to the spread of undetected defects, cracks or other similar damage. In the manufacture of structures in which any damage is allowed, two main problems must be solved. These problems consist in ensuring a controlled safe growth of defects, i.e., safe operation with cracks, and in compulsory containment of damage, as a result of which either residual durability or residual strength should be provided. In addition, the calculation of the permissible damage does not exclude the need for a thorough analysis and calculation of fatigue.

The main point which is based on the concept of safe damage consists in the fact that there are always defects, even in new designs, and that they may remain undetected. Thus, the first condition for the admissibility of the defect is a condition that any element of the design, including all additional units to transmit the load, must allow safe operation in the presence of cracks.

Control of the safe growth of defects. The occurrence of fatigue cracks can be avoided by creating a structure at all points where the stresses are below a certain level. However, lowering the stress level leads to an increase in the weight of the structures. In addition, cracks can occur not only from fatigue, but also for other reasons, for example, due to accidental damage obtained during operation, or due to material defects. Therefore, in real construction, a number of small cracks in the structure at the time of release from the plant are allowed. Large of these cracks can develop during operation.

The most important element of the safety principles of damage becomes a time period during which the crack can be detected. Due to various contingencies probability of detecting cracks when viewed unstable. Sometimes scarcely visible cracks found in the most remote zones of the construction and at the same time can be passed very largecracks elsewhere. For the case when was missed during the inspection "Boeing-747» crack length 1800 mm below the fairing in the pressurized cabin of an aircraft.

Therefore, for the structural elements determining the carrying capacity of the airframe, a program for monitoring the destruction must be compiled. An important element of the destruction control program is the development of verification methods. For each element, appropriate verification methods should be developed and proposed. For individual parts of the elements, non-destructive methods of monitoring different sensitivity may be required. The timing of the verification is established on the basis of an analysis of available information on the growth of the crack, taking into account the initial size of the defect and the size of the detected defect, which depends on the sensitivity of the flaw detection method used. The timing of the verification should be based on ensuring that, provided that the required safety factor is achieved, the undetected defect does not reach a critical size until the next test. Usually the intervals between the next checks are assigned so that two checks are made before reaching any crack of a critical size.

Safety principles of damage to aircraft design necessitated greater use of non-destructive methods for monitoring the technical condition of functional systems. The possibilities of different methods of non-destructive testing for the detection of fatigue cracks. NDT methods are constantly being improved.

Fatigue, corrosion and crack resistance. In the practice of the operation of the aircraft, numerous cases of destruction of the details of elements and assemblies from the fatigue of the material are known. Such destruction is the result of the action of variables or repeated loads. And for fatigue failure requires a much smaller maximum load than with static destruction. In flight and while moving on the ground, many components and structural elements of the aircraft are subjected to variable loads and although nominal voltages are often low, stress concentration, which usually does not reduce the static strength, can lead to fatiguedestruction. This is confirmed by the practice of exploitation of not only the sun but also ground vehicles. Indeed, you can almost always observe the fatigue failure and very rare - the destruction of the static loads.

A feature of fatigue failure is the absence of deformations in the fracture zone. Similar phenomena are observed even in materials such as soft steels that are highly plastic in the case of static destruction. This is a dangerous feature of fatigue failure, since there are no signs preceding destruction. Nascent fatigue symptoms are usually very small and difficult to detect until they reach a macroscopic size. Further they spread rapidly and in a short period of time complete destruction occurs. Thus, the timely detection of fatigue cracks is a difficult task. The most common fatigue cracks are generated in the zone of shape change or surface defects of parts.

Such defects, as well as a small change in the working section of the details do not affect the static strength, as plastic deformation reduces the effect of stress concentration. At the same time, fatigue failure of parts plastic deformations tend to be small, thereby reducing the concentration of stress in the area and there is no account of the concentrationstress is essential, however, it is important in the design of components operating under varying loads, making them easier and safer against fatigue failure.

Thus, the factors influencing the fatigue resistance include: stress concentrators, dimensions of parts, the relative importance of both static and cyclic loads and corrosion, especially corrosion of friction, which is the result of repeated small movements of the two contacting surfaces.

Fatigue is usually caused by the destruction of many thousands or millions of load cycles. However, they can occur after tens or even hundreds of cycles.

All elements, parts and components of the aircraft are subject to dynamic loads when moving on land and in flight. Variable loads of various types, acting on structural elements, parts of aggregates and devices, cause the corresponding variable stresses, which ultimately lead to fatigue failure. The speed of mechanical destruction processes of loaded parts and assemblies, respectively, and the time to failure depend on the structure and properties of materials, on the stresses caused by the operating loads, on temperature and other factors. However, the nature of failure from fatigue of the material has a peculiar form, different from brittle fracture.

Fatigue failure of parts usually begins near the metallurgical or technological defects, stress concentration zones, as well as the presence of technological defects in the products.

As is known, static failure is mainly determined by the probability of occurrence of a large load in flight, for example, by a gust of air as a result of which the Sun will operate the load exceeds the limits of the static strength of the structure, ie, the possibility of static destruction - it is essentially a question of probability of occurrence of a large load.

Fatigue failure under these assumptions - the result of the application of a sufficient number of load cycles, or a sufficient number of flights Sun at a certain distance.

The main difference between fatigue and static loading is as follows:

a major factor in fatigue strength for a given distribution of loads, even with the scatter of the data is the number of load changes or service life; for the static strength and the destruction - of the load;

The nature of the probabilistic approach to fatigue loading differs significantly from the nature of the probabilistic approach to static loading-for specific operating conditions the probability of the effect of a single large load on the aircraft, for example, from an air gust exceeding static destructive, does not depend on the time of operation. This can occur at the beginning and end of life. The likelihood of fatigue failure varies during operation, significantly increasing at the end of life. At the same time, designers and scientists believe that the designated resource or life-span and the corresponding probability level should be such that the frequency of recurrence of destruction has a sufficiently small value, which, if possible, would be generally accepted. This probability value is 10 9, which is taken as the basis for leading foreign and domestic aviation companies.

Aviation experts believe that the corrosion fatigue as well as damage to the same extent determines the service life of the aircraft structure. Most sources of corrosion - structural damage when loading the sun on the ground and scratched skin.

It is known that corrosion damage to the structure is entirely dependent on the operating conditions and the quality of the Armed Forces service.

The instructions, first of all, attention is drawn to the corrosion of the main structural elements of power. It is found that the corrosion is caused by a more internal than external factors. Thus, the cause of corrosion - liquid spilled in the area of ​​the buffet (especially fruit juice) and toilets.

Areas of the fuselage structure, are particularly susceptible to corrosion and fatigue cracks (shaded).

The least dangerous in relation to the total fatigue (uniform) corrosion. But in actual use uniform corrosion in its pure form is rare and is usually supplemented with ulcerative lesions. The effect of such corrosion fatigue resistance.

It can be seen that depending on the area and depth of corrosion damage, fatigue life of an alloy D16T significantly reduced. The area of ​​corrosion damage reduces the fatigue resistance of less than the diameter and depth of corrosion pits.

When using the process of accumulation of fatigue and corrosion damage alternate with partial overlapping each other. It is usually assumed that the corrosive lesions develop on parking, and fatigue - in flight. Corrosion damage is stress concentrators.

Terms and approaches used in the justification of resources within 103 l. h for 20-25 years of operation, determine the need to use while ensuring safety at the present stage, along with the principle of "safe-life" as a progressive principle of "Safe damage."

This last principle allows fatigue damage to the structural elements during the time interval between two consecutive inspections under the conditions that the interval is not too great, damage does not reach its limiting state, and does not lead to destruction of the structure as a whole.

Consequently, the strength criterion of the aircraft, claiming the inadmissibility of cracking, incorrect for the structure as a whole, as in a long-term operation of aircraft virtually impossible to avoid fatigue cracks in some of its elements. It is necessary to find a crack in time and prevent their further development for the maximum allowable size.

Thus, the strength resource of the aircraft should be based on the criterion of strength, taking into account the intensity of the origin and development of cracks for design in general, and in the elements that do not lead to a catastrophic outcome.

There is the concept on which it is believed that during the 30 min. 101 l. h should be safe, and then to 60 * 103 l. h - operation provided by the structural properties of survivability.

Recall that under the sun vitality or functional systems refers to the property that provides the proper performance of the functions specified in the flight (or flights) with individual faults or damage their elements or nodes. It is ensured by the provision, the specific design solutions, favoring rather slow development of damage and sufficient strength in the presence of a fault to be readily available for the detection of damage and objective control, if possible.

Experience shows that during prolonged wear of operation, fatigue and corrosion damage are the most massive failures.

Fatigue cracks lead to a decrease in strength of the structure and determine its strength reliability. Therefore, the design must be provided that the following conditions: the development and distribution of cracks in structural elements should be so slow that the residual static strength in the development of cracks to the size of its visual detection was sufficient for trouble-free operation of the sun without restrictions.

Let us consider some of the results of testing the samples of the fuselage shell of an aircraft with an airtight cabin. So, the scheme of development of a fatigue crack in the fuselage panels of a DC-10 aircraft is shown. The residual strength of the fuselage of the DC-10 aircraft was investigated on panels of 4267 x 2642 mm size with a radius of curvature ZOU mm. The tests were carried out under conditions of combined loading simulating inertial loads and boost pressure in the passenger cabin. To do this, take the panel from the top of the skin with the existing initial crack equal to 12 mm. As can be seen, during the first stage of testing at a nominal pressure of 0,65 Pa up to 15 000 cycles, crack growth was practically not observed. After making the notch in the power element and some increase in internal pressure, the crack growth rate began to increase, not reaching, however, a dangerous value. At 46 000 cycles, the central frame collapsed, then the two frames were destroyed, which resulted in a sharp increase in the rate of crack development and destruction of other force elements. The complete destruction of the panel occurred at a crack length of 1157 mm and at a pressure exceeding in 1,53 times the nominal pressure in the cabin.

Similar tests conducted on other panels with a set of security elements, have shown the ability to create designs of increased vitality and of the principle of "safe" of damage to the structure ensuring monitoring its condition at the MOT.

However, the most dangerous fatigue failure of structural elements of the fuselage. For example, cracks in the skin of the fuselage of the aircraft "Comet", appeared near the cutouts for windows, caused the two accidents of this type of aircraft.

The main reason for re-cracking load of the fuselage skin with pressurized cockpit aircraft "Comet" and design flaws. As is known, the aircraft skin undergoes repeated tension-compression load. They led to the development of cracks in the stress concentration. After performing plating crack completions of this type were not observed.

The design allows for increased survivability of certain dimensions of the damage that must meet the more general regulatory requirements. For example, the company "Douglas" believes that the residual strength of the structure of the passenger aircraft must be provided at the fracture wing length 400 mm disrupted middle stringer and in the fuselage for longitudinal crack length 1000 mm disrupted middle titanium stopper or transverse crack up to 400 mm destroyed the middle spar.

The company "Lockheed" determines the following possible damage to the fuselage: a crack in the skin may be long 300 mm destroyed in the middle of the frames or stringer; longitudinal crack in the skin - up to 500 mm; crack, running from the corner of a cut-out to 300 mm with the destruction of the frame or stringer.

The ICAO requirements specified that a minimum level of residual strength of the damaged structures must match the maximum operating load of 66,6% estimated for the calculation of the most important cases of loading.

GOST 27.002 83 defines durability as a property of the object continue to operate until a certain status in the installed system AMO. The limit condition can be caused by: fatal violation of safety requirements due to violation of the structural strength; unavoidable care units for the parameters of tolerance; unavoidable reduction in effectiveness; the need to perform major repairs in accordance with current regulatory and technical documentation.

Like reliability, durability is laid in the design of the aircraft, it is provided in the production and maintained in the process of operation. For AT, longevity is determined from the condition of flight safety and the desirability of its further application on the basis of comparative efficiency and the possibility of replacing with better samples. When designing AT products take into account possible loads during operation, operating modes; Select the appropriate material for the parts, methods of technological processing. For elements operating under friction conditions, the materials most wear-resistant in the intended operating conditions are selected, and so on.

All this allows designers to not only create a workable design, but also to carry out the relevant calculations can ensure the required standards of durability designed equipment.

Durability as a property of the structure depends on numerous factors, which can be divided into strength, operational and organizational.

Strength include design, manufacturing, processing, load and temperature factors. Among them are stress concentrators in the elements of construction and residual stresses resulting from the imperfect technology and due to plastic deformation in the assembly of parts and repairs; properties of materials and their change during the operation, including an initial static strength; fatigue limit; the stress intensity factor for the type of separation and destruction of the shift.

Experts believe that using modern achievements of science, engineering and technology, we can ensure the longevity of the structure of the main parts of the aircraft to 40 • 103 l. h. Without cracking aircraft can bump 30 103 x x l. h. If we assume that the cost-effective life (or duration of operation) is 60 • 103 l. h, it is possible to provide a guaranteed about half of this period, the sun and the other half will be operated with damage tolerance parts and assemblies and their replacement during repairs.